To investigate the crashworthiness and energy absorption characteristics of aircraft fuselage substructures and to conduct structural crashworthiness design, this study focuses on a typical metal aircraft fuselage as the research object. A drop test of a typical fuselage substructure was performed, and the energy absorption characteristics were evaluated based on both experimental and simulation analysis results. Subsequently, an energy absorption design for the fuselage substructure was developed. The influence of structural layout parameters on the crash response of the new fuselage substructure was examined through simulation analysis. Comparisons were made regarding structural deformation, impact force-time curves, acceleration responses, and energy absorption for fuselage substructures with different layout parameters. The results indicate that during the crash process of the original structure, the primary energy absorption modes include plastic deformation and fracture in the column, frame, and beam connection areas, bending deformation of cabin floor beams, and failure of connectors. Since all columns bend and break near the connection areas, the other parts of the columns are almost entirely free of plastic deformation, resulting in limited energy absorption by the columns. The new substructure proposed in this study features a novel configuration that fully leverages the advantages of metal plastic deformation. Compared to the original configuration, the new substructure exhibits more uniform deformation while maintaining the same total mass of the fuselage structure. It significantly reduces the peak load and acceleration at the early stage of the crash. The proportion of energy absorption by the frame and energy-absorbing components has increased markedly. After optimization, the average overload of the new fuselage substructure is reduced by 30.8% compared to the original configuration. The average acceleration of the two mass points on the cabin floor of the new fuselage substructure is reduced by 25.0% and 37.6%, respectively, compared to the original configuration. These findings provide valuable insights and references for the crashworthiness design of aircraft fuselage substructures.

The small-scale test has several advantages, such as low cost, low risk, and short duration, and has been widely applied in aerospace and other fields. Taking the lower structure of a typical civil aircraft fuselage as the research object, this study conducted theoretical analysis and experimental methodology of scaling on the impact crashworthiness of civil aircraft structures. Using a dimensional analysis, the complex dynamics of the fuselage crash were simplified to identify key physical parameters and processes. The main objects, critical physical parameters, and physical processes involved in the aircraft crash were discussed, leading to the extraction of key basic physical parameters and the derivation of primary dimensionless numbers that control the crash response of the fuselage structure. Based on the Buckingham Π theorem, the scaling factor for civil aircraft crashes was derived, establishing the small-scale experimental methodology. A 1/4 scale experimental model was designed and fabricated, and an impact test at a speed of 6 m/s was performed. The velocity, acceleration, ground impact load, deformation, and failure modes of key components in both full-scale and small-scale crash tests were obtained and compared. The applicability and accuracy of the small-scale theory in the crash experiment of the civil aircraft fuselage frame section were verified. The results show that the deformation and failure modes of the frames and columns of the 1/4 scale model are in good agreement with those of the full-scale model. The peak crash load prediction error of the small-scale structure for the full-scale prototype structure is 14.4%, the peak seat acceleration prediction error is 14.8%, and the peak acceleration prediction error at the beam is 13.1%. The small-scale tests can effectively predict the deformation, failure process, and dynamic response of key parts of the full-scale prototype structure. Therefore, the small-scale test could be used to verify and evaluate the crash performance of civil aircraft structures.

Addressing the frequent occurrence of landing gear shimmy under maneuvering conditions, this paper conducts research on high-fidelity modeling and parameter coupling effects. Based on a delay tire dynamic model, the coupled landing gear—tire shimmy dynamics equations are developed, which incorporate the time-varying tire lateral deformation modes. Utilizing Hopf bifurcation theory, we compute the global dynamic characteristics of the system through multi-dimensional projections, quantitatively revealing the influence patterns and mechanisms of key parameters (such as lateral bending stiffness and torsional stiffness) on the stability of the landing gear system across various taxiing speeds. Comparative results with traditional tire models show consistency of both and limitations of traditional tire models in capturing parameter coupling effects. By employing an enhanced Bayesian optimization method, we optimize the bifurcation characteristics of the landing gear system, resulting in a significant improvement in stability. Based on the parameter optimization results and combined with Sobol global sensitivity analysis, the comparative effects of multi-dimensional parameters on system stability are quantitatively revealed, with the results complementing and validating those from the Hopf bifurcation analysis. The study identifies that when the torsional mode frequency of the landing gear approaches the lateral bending mode frequency, substantial modal coupling occurs, creating a bistable shimmy critical boundary that drastically reduces stability. Low lateral stiffness of the landing gear may trigger shimmy during low-speed taxiing, while greater lateral stiffness at high speeds necessitates matched damping enhancement for shimmy mitigation. Increasing torsional stiffness and decreasing tire lateral stiffness yield notable shimmy suppression effects across the entire speed spectrum, indicating their potential as key control parameters for targeted design and optimization.

High-energy laser weapons have gradually emerged as a critical new destructive capability on modern battlefields, complementing traditional kinetic weapons. In real service conditions of aircraft, external structures such as skins are subjected to long-term cyclic loads due to air pressure. Therefore, it is essential to consider the influence of boundary loads when studying laser-induced damage. Focusing on the problem of high-energy laser damage effects under boundary loading conditions, this study conducted experiments on a typical 2A12 aerospace aluminum alloy under high-energy continuous laser irradiation. By integrating a boundary preload loading apparatus, the effects of boundary preloads on material response and damage behavior were analyzed. It was found that the primary damage mode under continuous high-energy laser irradiation involves molten material flowing downward under gravitational forces to form perforations. Under boundary preload conditions, thermo-expansion and thermal softening jointly influence the thermomechanical response of the target. Observations of boundary load curves during the ablation process revealed three distinct stages. After laser irradiation ceased, residual boundary loads in tensile and compressive preloading conditions recovered to approximately 80% and 25%, respectively, indicating irreversible damage akin to plastic deformation. Furthermore, a thermomechanical coupling simulation model incorporating phase transformation and material flow was established based on experimental results. By analyzing the degradation pattens of thermal expansion coefficients and elastic modulus, the roles of thermo-expansion and thermal softening during laser ablation were elucidated.

The full-aircraft landing ground load is one of the basic inputs for transport aircraft design and performance optimization. In order to study the characteristics of the ground load and its influencing factors during the full-aircraft landing, we take the typical single-channel simplified aircraft test piece as the object, designs the full-aircraft landing test scheme, and establishes the full-aircraft landing simulation analysis model which considers the flexibility of the fuselage. Through experimental and simulation analysis, the influence laws of landing conditions, fuselage weight, and fuselage flexibility on its ground load are discussed. The results show that the aircraft ground load obtained from the full-aircraft landing experiment can effectively consider influence of fuselage flexibility. The full-aircraft landing simulation model has a good prediction accuracy, the error between the curve peak and the ground load experimental result is within 5%, and the correlation coefficient is greater than 0.967. The landing velocity has the most significant effect on the aircraft landing load, and the sensitivity coefficient is 1.084. The effect of the pitch angle is the weakest, and the sensitivity coefficient is only 0.030. With the increase of landing weight, the ground load and the energy absorption amount of the buffer increase significantly, the buffer efficiency gradually decreases, and the sensitivity of the ground load peak, the buffer absorption energy, and buffer efficiency to the landing weight are 0.918, 1.131, and 0.086, respectively. With the decrease of the stiffness of the fuselage, the flexibility effect of the fuselage is significantly enhanced, and the ground load peak, the buffer absorption energy, and the efficiency gradually decrease. The influence of the fuselage flexibility on the transport aircraft landing load cannot be ignored.

The crashworthiness of an aircraft is the focus of its structural safety design and evaluation, and the relationship between the fuselage frame segment and the full-aircraft structure in terms of crash response has been one of the unresolved issues. In this paper, a vertical drop test with a fuselage frame segment with cabin door and a typical full-scale civil aircraft were conducted, and the results were compared. Then, a dynamic model of the full-aircraft crash was established and validated based on the experimental data from the full-aircraft crash. Furthermore, the model trimming method was used to obtain the crash dynamics models of the fuselage frame segment with a cabin door, the forward fuselage, and aft fuselage, and simulations were conducted under the experimental conditions. Based on the simulation results, differences between the crash responses of different fuselage frame segments and the full-aircraft structure were explored. The results show that the established full-aircraft crash dynamics model has high predictive accuracy, the predicted dynamic response of the structure is consistent with the experimental, and the deformation and movement process of the structure are consistent the experimental results, proving the accuracy of the model trimming method. The deformation of the floor structure and the floor acceleration of the fuselage frame segment with a cabin and the aft fuselage are less than those of the full-aircraft structure, the maximum deviation of the acceleration peak is 13.6%, and the damage modes are significantly different. For forward fuselage, the deformation is greater than that of the full-aircraft structure, but the floor acceleration is still less than that of the full-aircraft structure, with a maximum deviation of the acceleration peak reaching 17.2%. This reveals that the boundary conditions and energy input of the full-aircraft and the fuselage frame segment experiment pieces are fundamentally different, leading to significant differences in crash response. the verification and evaluation of aircraft crashworthiness, different verification methods and objects should be used for different verification targets. Using the full-aircraft structure for crashworthiness verification provides the most realistic results, while using the fuselage frame segment can verify the crash simulation model to support the simulation-based aircraft crashworthiness assessment. Additionally, fuselage segment with a cabin door and the aft fuselage are greatly affected by the contraction of the aircraft's nose and tail during a crash, and thus are not suitable as verification objects for aircraft crashworthiness. When using the forward fuselage as the verification object for aircraft crashworthiness, the end frame of the experiment needs to be strengthened to simulate actual boundary stiffness conditions as closely as possible.

The environmental factors in survivable civil aircraft crash accidents have a significant impact on passenger evacuation. Based on the characteristics of typical survivable civil aircraft crash accident scenarios, this paper proposes an experimental method for passenger emergency evacuation that can simulate typical crash accident environments, and introduces simulation methods for aircraft body attitudes, environmental sound/light, cabin obstacles, visual environment of portholes, and high-precision measurement methods for the personnel evacuation trajectory. The impact of crash environment factors such as aircraft body tilt attitudes, accident sound/light environments, and aisle obstacles on the emergency evacuation of civil aircraft passengers is studied through volunteer experiments. The results indicate that the proposed experimental method for emergency evacuation considering the impact of crash environment and the simulation scheme for crash environment are feasible, and obtain effective experimental data; the aircraft body attitude change within a pitch angle of ± 5° has no significant impact on the total evacuation time and evacuation efficiency of multi-passenger evacuation, and the difference in the speed of passengers moving uphill and downhill in the aisle area increases with the increase of pitch angle; the sound and lighting environment of the crash accident have a certain impact on the emergency evacuation of passengers, and the accident background sound can reduce the total evacuation time, while emergency lighting increases the total evacuation time; obstacles in the cabin aisle overall increase the total evacuation time, with a shorter distance between the obstacles and the exit resulting in lower evacuation efficiency; compared to movable obstacles, fixed obstacles can lead to lower evacuation efficiency.

The crash test of civil aircraft is a worldwide technical problem, and it is the most direct means to evaluate the crashworthiness of civil aircraft. In this paper, the high precision lifting height control and high reliability delivery method of the full-scale aircraft crash test are proposed, and the testing methods of the key physical parameters such as structural response and dummy response are given. A dynamic response test system for the whole aircraft crash test was constructed, and using a unified time reference trigger method, the ground impact load, structural acceleration response, dummy response, and aircraft failure and deformation were analyzed. The response distribution rules of different parts of the aircraft were obtained. The revised comprehensive evaluation index ICI of the adaptability was put forward. The results show that the test data are complete and reliable. After the vertical crash at 5.71 m/s, the lower structure of the cabin floor is seriously deformed, and the upper structure of the fuselage in the central wing area is obviously deformed due to the inertia effect of the wing. The stiffness difference of the different fuselage segments results in significant differences in crash load and dynamic response. The higher the stiffness is, the smaller the deformation and the greater the acceleration response will be. After the crash, the load on the passengers was within safe range, the cabin seats are intact, the cabin doors can be opened normally., the living space of the passengers is sufficient, and the evacuation channel of the passengers is unblocked. Compared with the original ICI index, the revised evaluation result has better engineering applicability.